1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to an air cooled large turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section in which a hot gas flow passes through and reacts against a plurality of stages of stationary guide vanes and rotary blades to drive a rotor shaft. The engine efficiency can be increased by providing for a higher temperature flow through the turbine. Modern blade and vane materials limit the temperature that can be used without damaging the airfoils. In order to increase the efficiency, some of the stages of vanes and blades are cooled by passing cooling air through the internal airfoils. This will allow for a higher operating temperature without damaging the airfoils. Complex cooling circuits have been proposed in the Prior Art to maximize the use of cooling air, since the cooling air is bled off from the compressor which also decreases the efficiency of the engine.
In large turbines such as an industrial gas turbine engine, the third stage rotor blade is very large, especially compared to aero engines. If cooling of the third stage rotor blade is required, cooling passages must be cast into the blade or drilled after casting.
Prior Art cooling of a large turbine rotor is achieved by drilling radial holes into the blade from blade tip and root sections. Limitations of drilling a long radial hole from both ends of the airfoil increases for a large highly twisted and tapered blade airfoil that are used in industrial gas turbine (IGT) engines. Reduction of the available airfoil cross sectional area for drilling radial holes is a function of the blade twist and taper. Higher airfoil twist and taper yield a lower available cross sectional area for drilling radial cooling holes. Cooling of the large, highly twisted and tapered blade by the prior art manufacturing technique will not achieve the optimum blade cooling effectiveness. Especially lacking cooling for the airfoil leading and trailing edges. This prevents high firing temperature applications as well as low cooling flow design. U.S. Pat. No. 6,910,843 B2 issued to Tomberg on Jun. 28, 2005 and entitled TURBINE BUCKET AIRFOIL COOLING HOLE LOCATION, STYLE AND CONFIGURATION discloses a large turbine blade (also referred to as a bucket) with radial cooling holes drilled into the blade.
U.S. Pat. No. 6,164,913 issued to Reddy on Dec. 26, 2000 and entitled DUST RESISTANT AIRFOIL COOLING shows a turbine airfoil with an internal cooling circuit having a triple-pass (3-pass) serpentine cooling circuit with a first leg adjacent to the airfoil leading edge, a second leg at mid-blade, and the third leg near the trailing edge and connected to exit holes on the trailing edge by metering holes. The 3-pass serpentine flow cooling circuit provides better cooling than the single pass straight radial holes of the Tomberg patent using the same amount of cooling flow because of the serpentine path through the blade.
It is an object of the present invention to provide for a cooling circuit within a large highly tapered blade.
It is another object of the present invention to provide a ceramic core assembly than can be used for casting a large highly tapered blade with a serpentine flow cooling circuit within the blade.